Deflector for gas turbine engine combustors and method of using the same

ABSTRACT

A deflector for a gas turbine engine combustor. The combustor includes a liner defining a combustion zone and a mixer assembly configured to supply the combustion zone with a predetermined mixture of fuel and air. The deflector includes a deflector body configured to couple to the liner. The deflector body includes a first surface configured to reflect thermal radiation to a predetermined focal area, and an aperture extending through the deflector body and configured to receive the mixer assembly therethrough.

BACKGROUND

The field of the disclosure relates generally to gas turbine enginesand, more particularly, to deflectors for use in gas turbine enginecombustors.

Combustors are used to ignite fuel and air mixtures in gas turbineengines to generate high energy working gases. Known combustors includean outer liner and an inner liner defining an annular combustion chamberin which the fuel and air are mixed and burned. A dome mounted at theupstream end of the combustion chamber includes mixers for mixing thefuel and air. Ignitors mounted downstream from the mixers ignite themixture, and it burns in the combustion chamber. At least somecombustors further include a deflector coupled to the dome andsurrounding the mixer that prevent hot combustion gases produced withinthe combustion chamber from impinging directly upon the dome andupstream components.

Air pollution concerns have led to stricter combustion emissionsstandards. These standards regulate the emission of nitrogen oxides(NOx), as well as other types of exhaust emissions, from the gas turbineengine. Generally, NOx is formed during the combustion process due tohigh flame temperatures in the combustor. At least some combustorsreduce NOx by using a lean fuel and air mixture that tends to lower theflame temperature. For lean mixtures, the volume percent of air in themixture may typically be increased up to a limit, sometimes referred toas a blowout limit, at which the air and fuel mixture can no longermaintain a flame. As such, to reduce NOx emissions, combustors typicallyuse a lean fuel and air mixture as close to the blowout limit aspossible. Lean fuel and air mixtures, however, may increase combustiondynamics and flame instability because of the low temperature flame.Flame instability and combustion dynamics generally reduces combustorand overall gas turbine engine efficiency and durability.

BRIEF DESCRIPTION

In one embodiment, a deflector for a gas turbine engine combustor isprovided. The combustor includes a liner defining a combustion zone anda mixer assembly that is configured to supply the combustion zone with apredetermined mixture of fuel and air. The deflector includes adeflector body that is configured to couple to the liner and including afirst surface that is configured to reflect thermal radiation to apredetermined focal area. The deflector body defines an apertureextending therethrough. The aperture is configured to receive the mixerassembly therethrough.

In another embodiment, a combustor for a gas turbine engine is provided.The combustor includes a liner defining a combustion zone. A mixerassembly is configured to supply the combustion zone with apredetermined mixture of fuel and air. The combustor further includes adeflector coupled to the liner. The deflector includes a deflector bodythat is configured to couple to the liner and including a first surfacethat is configured to reflect thermal radiation to a predetermined focalarea. The deflector body defines an aperture extending therethrough. Theaperture is configured to receive the mixer assembly therethrough.

In a further embodiment, a method for stabilizing a flame within a gasturbine engine combustor is provided. The combustor includes a linerdefining a combustion zone, a mixer assembly, and a deflector coupled tothe liner. The deflector includes a deflector body that is configured tocouple to the liner. The deflector body includes a first surface anddefines an aperture extending therethrough. The aperture is configuredto receive the mixer assembly therethrough. The method includesgenerating the flame within the combustion zone using a predeterminedmixture of fuel and air supplied by the mixture assembly. The methodfurther includes reflecting thermal radiation of the flame to apredetermined focal area.

DRAWINGS

These and other features, aspects, and advantages of the presentdisclosure will become better understood when the following detaileddescription is read with reference to the accompanying drawings in whichlike characters represent like parts throughout the drawings, wherein:

FIG. 1 is a schematic, cross-sectional illustration of an exemplaryturbofan engine in accordance with an example embodiment of the presentdisclosure.

FIG. 2 is a cross-sectional view of an exemplary combustor that may beused with the turbofan engine shown in FIG. 1.

FIG. 3 is a perspective view of an exemplary deflector that may be usedwith the combustor shown in FIG. 2.

Unless otherwise indicated, the drawings provided herein are meant toillustrate features of embodiments of this disclosure. These featuresare believed to be applicable in a wide variety of systems comprisingone or more embodiments of this disclosure. As such, the drawings arenot meant to include all conventional features known by those ofordinary skill in the art to be required for the practice of theembodiments disclosed herein.

DETAILED DESCRIPTION

In the following specification and claims, reference will be made to anumber of terms, which shall be defined to have the following meanings.

The singular forms “a,” “an,” and “the” include plural references unlessthe context clearly dictates otherwise.

“Optional” or “optionally” means that the subsequently described eventor circumstance may or may not occur, and that the description includesinstances where the event occurs and instances where it does not.

Approximating language, as used herein throughout the specification andclaims, may be applied to modify any quantitative representation thatcould permissibly vary without resulting in a change in the basicfunction to which it is related. Accordingly, a value modified by a termor terms, such as “about,” “approximately,” and “substantially,” are notto be limited to the precise value specified. In at least someinstances, the approximating language may correspond to the precision ofan instrument for measuring the value. Here and throughout thespecification and claims, range limitations may be combined and/orinterchanged; such ranges are identified and include all the sub-rangescontained therein unless context or language indicates otherwise.

Embodiments of a deflector for a gas turbine engine combustor asdescribed herein provide an apparatus for lean combustion systems thatfacilitates increasing combustor flame stability and reducing emissionsthereof. Specifically, the deflector includes a body with a downstreamsurface that reflects thermal radiation of a combustor flame within thecombustor to a predetermined focal area. The downstream surface isformed in a predetermined shape, such as but not limited to, a parabolicshape to facilitate thermal radiation reflection from the downstreamsurface to the predetermined focal area. Furthermore, the deflector mayalso include a reflective thermal barrier coating that furtherfacilitates reflecting thermal radiation of the combustion flame. Byreflecting thermal radiation from the deflector and back into thecombustion flame, reactants and products within cooler areas of thecombustion flame may increase in temperature to facilitate stabilizingthe combustion flame while maintaining a lean fuel/air mixture.Stabilizing the combustion flame reduces combustion dynamics and allowsfor the combustor to operate closer to a lean blowout limit to reduceNOx emissions.

FIG. 1 is a schematic cross-sectional view of a gas turbine engine inaccordance with an exemplary embodiment of the present disclosure. Inthe exemplary embodiment, the gas turbine engine is a high-bypassturbofan jet engine 110, referred to herein as “turbofan engine 110.” Asshown in FIG. 1, turbofan engine 110 defines an axial direction A(extending parallel to a longitudinal centerline 112 provided forreference) and a radial direction R (extending perpendicular tolongitudinal centerline 112). In general, turbofan engine 110 includes afan case assembly 114 and a gas turbine engine 116 disposed downstreamfrom fan case assembly 114.

Gas turbine engine 116 includes a substantially tubular outer casing 118that defines an annular inlet 120. Outer casing 118 encases, in serialflow relationship, a compressor section including a booster or lowpressure (LP) compressor 122 and a high pressure (HP) compressor 124; anannular combustion section 126 including a plurality ofcircumferentially spaced fuel nozzle assemblies 218 (shown in FIG. 2); aturbine section including a high pressure (HP) turbine 128 and a lowpressure (LP) turbine 130; and a jet exhaust nozzle section 132. A highpressure (HP) shaft or spool 134 drivingly connects HP turbine 128 to HPcompressor 124. A low pressure (LP) shaft or spool 136 drivinglyconnects LP turbine 130 to LP compressor 122. The compressor section,combustion section 126, turbine section, and exhaust nozzle section 132together define an air flow path 138.

In the exemplary embodiment, fan case assembly 114 includes a fan 140having a plurality of fan blades 142 coupled to a disk 144 in a spacedapart manner. As depicted, fan blades 142 extend outwardly from disk 144generally along radial direction R. Fan blades 142 and disk 144 aretogether rotatable about longitudinal centerline 112 by LP shaft 136.

Referring still to the exemplary embodiment of FIG. 1, disk 144 iscovered by a rotatable front hub 146 aerodynamically contoured topromote an airflow through plurality of fan blades 142. Additionally,exemplary fan case assembly 114 includes an annular fan casing or outernacelle 150 that circumferentially surrounds fan 140 and/or at least aportion of gas turbine engine 116. Nacelle 150 is supported relative togas turbine engine 116 by an outlet guide vane (OGV) assembly 152.Moreover, a downstream section 154 of nacelle 150 may extend over anouter portion of gas turbine engine 116 so as to define a bypass airflowduct 156 between nacelle 150 and outer casing 118.

During operation of turbofan engine 110, a volume of air 158 entersturbofan engine 110 through an associated inlet 160 of nacelle 150and/or fan case assembly 114. As air 158 passes across fan blades 142, afirst portion of air 158 as indicated by arrow 162, known as fan streamair flow, is directed or routed into bypass airflow duct 156 and asecond portion of air 158 as indicated by arrow 164 is directed orrouted into air flow path 138, or more specifically into boostercompressor 122. The ratio between first portion of air 162 and secondportion of air 164 is commonly known as a bypass ratio. The pressure ofsecond portion of air 164 is then increased, forming compressed air 166,as it is routed through booster compressor 122 and HP compressor 124 andinto combustion section 126, where it is mixed with fuel 168 and burnedto provide combustion gases 170.

Combustion gases 170 are routed through HP turbine 128 where a portionof thermal and/or kinetic energy from combustion gases 170 is extractedvia sequential stages of HP turbine stator vanes 172 that are coupled toouter casing 118 and HP turbine rotor blades 174 that are coupled to HPshaft or spool 134, thus causing HP shaft or spool 134 to rotate,thereby supporting operation of HP compressor 124. Combustion gases 170are then routed through LP turbine 130 where a second portion of thermaland kinetic energy is extracted from combustion gases 170 via sequentialstages of LP turbine stator vanes 176 that are coupled to outer casing118 and LP turbine rotor blades 178 that are coupled to LP shaft orspool 136, thus causing LP shaft or spool 136 to rotate, therebysupporting operation of booster compressor 122 and/or rotation of fan140. Combustion gases 170 are subsequently routed through jet exhaustnozzle section 132 of gas turbine engine 116 to provide propulsivethrust. HP turbine 128, LP turbine 130, and jet exhaust nozzle section132 at least partially define a hot gas path 180 for routing combustiongases 168 through gas turbine engine 116. Simultaneously, the pressureof fan stream air 162 is substantially increased as fan stream air 162is routed through bypass airflow duct 156, including through outletguide vane assembly 152 before it is exhausted from a fan nozzle exhaustsection 182 of turbofan engine 110, also providing propulsive thrust.

It should be appreciated, however, that the exemplary turbofan engine110 depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, turbofan engine 110 may have any other suitableconfiguration. It should also be appreciated, that in still otherexemplary embodiments, aspects of the present disclosure may beincorporated into any other suitable gas turbine engine. For example, inother exemplary embodiments, aspects of the present disclosure may beincorporated into, e.g., a turboprop engine, a military purpose engine,and a marine or land-based aero-derivative engine.

FIG. 2 is a cross-sectional view of an exemplary combustor 200 that maybe used with turbofan engine 110 (shown in FIG. 1). In the exemplaryembodiment, annular combustion section 126 includes a combustor 200having a combustion zone or chamber 202 defined by annular, radiallyouter and radially inner liners 204 and 206. Specifically, outer liner204 defines a radially outer boundary of combustion chamber 202, andinner liner 206 defines a radially inner boundary of combustion chamber202. Liners 204 and 206 are spaced radially inward from an annularcombustor casing 208 which extends circumferentially around liners 204and 206. Combustor 200 also includes an annular dome 210 mountedupstream from outer and inner liners 204 and 206 respectively. Dome 210defines an upstream end of combustion chamber 202 and a plurality ofmixer assemblies 212 are spaced circumferentially around dome 210 todeliver a mixture of fuel and air to combustion chamber 202. Each mixerassembly 212 includes a pilot mixer 214 and a main mixer 216. Main mixer216 is concentrically aligned with respect to pilot mixer 214 andextends circumferentially around pilot mixer 214. A plurality ofcircumferentially spaced and axially-extending fuel nozzle assemblies218 are coupled in flow communication with each respective mixerassembly 212. Furthermore, in the exemplary embodiment, combustor 200includes one or more deflectors 220 that are coupled to and spacedcircumferentially around dome 210 at each mixer assembly 212 location.Downstream of mixer assembly 212 and deflector 220 is an igniter 222that extends through outer casing 208 and into combustion chamber 202 toprovide initial ignition of the mixture of compressed air 166 and fuel168. In various embodiments, igniter 222 may provide continuous orintermittent ignition support to combustion chamber 202.

In operation, combustor 200 receives compressed air 166 discharged fromHP compressor 124 in a diffuser section 224 at flow upstream ofcombustion chamber 202. A portion of the flow of compressed air 166 ischanneled through mixer assembly 212. At mixer assembly 212 compressedair 166 is mixed with fuel 168 from fuel nozzle assembly 218 anddischarged into combustion chamber 202 where the mixture of air 166 andfuel 168 is ignited by igniter 222 creating a flame 224 withincombustion chamber 202 that burns the mixture and provides combustiongases 170 that are channeled downstream to HP turbine 128 (shown in FIG.1). In the exemplary embodiment, combustor 200 is a lean combustor.Specifically, at engine start conditions and engine low power operation,combustor 200 uses only fuel 168 provided to the pilot mixer 214 forgenerating combustion gases 170. At pilot mixer 214, fuel 168 includes apilot fuel stream 226 that is mixed with a first portion 228 ofcompressed air 166 to provide a rich mixture (higher fuel 226 to air 228ratios within the mixture) that is ignited for a pilot flame 230 withina region 232 that is adjacent to pilot mixer 214. At engine high poweroperation, combustor 200 uses fuel 168 split between pilot mixer 214 andmain mixer 216 for generating combustion gases 170. At main mixer 216,fuel 168 includes a main fuel stream 234 that is mixed with a secondportion 236 of compressed air 166 to provide a lean mixture (lower fuel234 to air 236 ratios within the fuel-air mixture) that is ignited for amain flame 238 within a region 240 that is adjacent to main mixer 216.At engine high power operation most of fuel 168 is injected through mainmixer 216 thus providing a lean burn combustion process to generatecombustion gases 170 while reducing NOx emissions.

Generally, pilot flame 230 burns at a higher temperature than main flame238 because the fuel 226 air 228 mixtures are richer. As such, duringengine start and engine low power operation combustion dynamics are lowleading to pilot flame 230/flame 224 that is stable. Main flame 238,however, generally burns at a lower temperature than pilot flame 230because the fuel 234 air 236 mixtures are leaner. As such, during highpower engine operation, when most of fuel 168 is injected through mainmixer 216, main flame 238/flame 224 instability may occur due to lowertemperatures leading to combustion dynamics. To facilitate reducing NOxemissions from the combustion process, flame 224 temperatures duringhigh power engine operation are reduced as low as possible and close toa blowout limit by increasing the air 236 to fuel 234 ratio because NOxis formed at high flame temperatures. As such, cooler temperatureregions, such as region 242, are formed within flame stabilization zonesthat decreases flame 224 stability and increases a likelihood of flameblowout.

In the exemplary embodiment, deflector 220 facilitates increasing flame224 stability and decreases combustion dynamics and a likelihood offlame blowout during engine operation at high power levels while usingthe lean burn combustion process. Specifically, deflector 220 reflects244 thermal radiation and infrared radiation generated by flame 224within combustion chamber 202 to a predetermined focal area 246 withincooler temperature region 242 to increase the temperature of region 242.Flame 224 generates thermal radiation by burning the fuel 168 and air166 mixture which typically heats up the surrounding combustorcomponents, such as outer and inner liners 204 and 206 and dome 210. Byreflecting 244 the thermal radiation back into flame 224, thetemperature of region 242 increases. The temperature of region 242increases by heating up entrained carbon dioxide and water vaportherein. Thus, reducing combustion dynamics and a likelihood of flameblowout, and increasing flame 224 stability.

It should be appreciated, that the exemplary combustor 200 illustratedin FIG. 2 is by way of example only, and that in other exemplaryembodiments, combustor 200 may have any other suitable configuration forlean based combustion. It should further be appreciated, that in stillother exemplary embodiments, aspects of the present disclosure may beincorporated into any other suitable gas turbine engine includingland-based aero-derivative engine combustion systems.

FIG. 3 is a perspective view of an exemplary deflector 220 that may beused with combustor 200 (shown in FIG. 2). In the exemplary embodiment,deflector 220 is substantially circular and includes a body 300 formedwith an aperture 302 sized to at least partially receive mixer assembly212 (shown in FIG. 2). Body 300 is coupled to dome 210 (shown in FIG.2), for example, by brazing. Body 300 includes an upstream surface 304and a downstream surface 306. Downstream surface 306 facilitatesreflecting thermal radiation from combustor flame 224 (shown in FIG. 2)to predetermined focal area 246 (shown in FIG. 2). For example,downstream surface 306 is parabolic such that downstream surface 306reflects thermal radiation from combustor flame 224 to predeterminedfocal area 246.

Parabolic shape 308 of deflector 220 receives thermal radiation on thecurved downstream surface 306 and reflects the thermal radiation topredetermined focal area 246. Parabolic curvature 308 of downstreamsurface 306 may be sized to position focal area 246 at any location thatfacilitates increasing temperature of combustor flame 224. Thus,reducing combustion dynamics and a likelihood of flame blowout, andincreasing flame 224 stability. In alternative embodiments, downstreamsurface 306 may have any other shape/contour that enables deflector 220to functions as described herein. Additionally, in alternativeembodiments, the substantially circular deflector body 300 may betrimmed to form a polygonal periphery.

In the exemplary embodiment, body 300 is fabricated from a superalloysubstrate 310 and coated with a thermal barrier coating 312 to reducethermal exposure when combustor 200 is operating. Physical vapordeposition thermal barrier coating, TBC 312, is applied to deflector 220and provides thermal protection thereto. Furthermore, TBC 312facilitates reflecting thermal radiation of flame 224 as describedabove.

During operation of combustor 200, deflector 220 protects dome 210 andmixer assembly 212 from hot gases and thermal flame radiation generatedwith combustion chamber 202. Furthermore, parabolic curvature 308 andTBC 312 reflect the thermal flame radiation back into flame 224 toincrease flame stability during lean combustor operation.

The above-described embodiments of a deflector for a gas turbine enginecombustor provide an apparatus for lean combustion systems thatfacilitates increasing combustor flame stability and reducing emissionsthereof. Specifically, the deflector includes a body with a downstreamsurface that reflects thermal radiation of a combustor flame within thecombustor to a predetermined focal area. The downstream surface isformed in a predetermined shape, such as, but not limited to, aparabolic shape to facilitate thermal radiation reflection from thedownstream surface to the predetermined focal area. Furthermore, thedeflector may also include a reflective thermal barrier coating thatfurther facilitates reflecting thermal radiation of the combustionflame. By reflecting thermal radiation from the deflector and back intothe combustion flame, reactants and products within cooler areas of thecombustion flame may increase in temperature to facilitate stabilizingthe combustion flame while maintaining a lean fuel/air mixture.Stabilizing the combustion flame reduces combustion dynamics and allowsfor the combustor to operate closer to a lean blowout limit to reduceNOx emissions.

An exemplary technical effect of the methods, systems, and apparatusdescribed herein includes at least one of: (a) reducing combustiondynamics in a gas turbine combustion system; (b) increasing leancombustion flame temperature; (c) increasing combustor flame stability;(d) reducing lean blowout limit of combustor; and (e) reducingcombustion emissions.

Exemplary embodiments of methods, systems, and apparatus for combustorflame stabilization are not limited to the specific embodimentsdescribed herein, but rather, components of the systems and/or steps ofthe methods may be utilized independently and separately from othercomponents and/or steps described herein. For example, the methods mayalso be used in combination with other systems requiring flamestabilization, and the associated methods, and are not limited topractice with only the systems and methods as described herein. Rather,the exemplary embodiment can be implemented and utilized in connectionwith many other applications, equipment, and systems that may benefitfrom thermal control.

Although specific features of various embodiments of the disclosure maybe shown in some drawings and not in others, this is for convenienceonly. In accordance with the principles of the disclosure, any featureof a drawing may be referenced and/or claimed in combination with anyfeature of any other drawing.

This written description uses examples to disclose the embodiments,including the best mode, and also to enable any person skilled in theart to practice the embodiments, including making and using any devicesor systems and performing any incorporated methods. The patentable scopeof the disclosure is defined by the claims, and may include otherexamples that occur to those skilled in the art. Such other examples areintended to be within the scope of the claims if they have structuralelements that do not differ from the literal language of the claims, orif they include equivalent structural elements with insubstantialdifferences from the literal language of the claims.

What is claimed is:
 1. A deflector for a gas turbine engine combustor,the combustor includes a liner defining a combustion zone and a mixerassembly configured to supply the combustion zone with a predeterminedmixture of fuel and air, said deflector comprising: a deflector bodyconfigured to couple to the liner and comprising a first surfaceconfigured to reflect thermal radiation to a predetermined focal area,wherein said deflector body defines an aperture extending therethrough,the aperture configured to receive the mixer assembly therethrough. 2.The deflector in accordance with claim 1, wherein said first surfacecomprises a parabolic shape.
 3. The deflector in accordance with claim1, wherein the predetermined focal area is positioned within thecombustion zone.
 4. The deflector in accordance with claim 1, whereinthe predetermined focal area is configured to heat relatively cool areasof the combustion zone to stabilize a flame during lean combustionwithin the combustor.
 5. The deflector in accordance with claim 1further comprising a thermal barrier coating formed on at least aportion of said deflector body.
 6. The deflector in accordance withclaim 5, wherein said thermal barrier coating is configured to reflectthermal radiation of a flame to the predetermined focal area.
 7. Thedeflector in accordance with claim 1, wherein said deflector bodycomprises a superalloy substrate configured to reduce thermal exposureto combustor components upstream of said deflector body.
 8. A combustorfor a gas turbine engine comprising: a liner defining a combustion zone;a mixer assembly configured to supply the combustion zone with apredetermined mixture of fuel and air; and a deflector coupled to saidliner and comprising: a deflector body configured to couple to the linerand comprising a first surface configured to reflect incident thermalradiation to a predetermined focal area, wherein said deflector bodydefines an aperture extending therethrough, the aperture configured toreceive said mixer assembly therethrough.
 9. The combustor in accordancewith claim 8, wherein said first surface comprises a parabolic shape.10. The combustor in accordance with claim 8, wherein said predeterminedfocal area is positioned within the combustion zone.
 11. The combustorin accordance with claim 8, wherein the predetermined focal area isconfigured to heat relatively cool areas of the combustion zone tostabilize a combustion flame during lean combustion within thecombustor.
 12. The combustor in accordance with claim 8 furthercomprising a thermal barrier coating formed on at least a portion ofsaid deflector body.
 13. The combustor in accordance with claim 12,wherein said thermal barrier coating is configured to reflect thermalradiation of a combustion flame to the predetermined focal area.
 14. Thecombustor in accordance with claim 8, wherein said deflector bodycomprises a superalloy substrate configured to reduce thermal exposureto combustor components upstream of said deflector.
 15. A method forstabilizing a flame within a gas turbine engine combustor, the combustorincluding a liner defining a combustion zone, a mixer assembly, and adeflector coupled to the liner, the deflector including a deflector bodyconfigured to couple to the liner, the deflector body including a firstsurface and defining an aperture extending therethrough, the apertureconfigured to receive the mixer assembly therethrough, said methodcomprising: generating the flame within the combustion zone using apredetermined mixture of fuel and air supplied by the mixture assembly;and reflecting thermal radiation of the flame to a predetermined focalarea.
 16. The method in accordance with claim 15, wherein the firstsurface has a parabolic shape.
 17. The method in accordance with claim16 further comprising reflecting thermal radiation into thepredetermined focal area using the parabolic shape of the first surface.18. The method in accordance with claim 15 further comprising formingthe first surface to focus the thermal radiation to the predeterminedfocal area positioned within the combustion zone.
 19. The method inaccordance with claim 15 further comprising forming a thermal barriercoating on at least a portion of the deflector body.
 20. The method inaccordance with claim 15 further comprising forming the body from asuperalloy substrate configured to reduce thermal exposure to combustorcomponents upstream of the deflector.